The present invention relates to gas turbine engines, and, in particular, relates to nozzles of gas turbine engines, and, more particularly, relates to the liners used in gas turbine engines.
It is well known in the jet engine art that extreme temperatures can develop in an exhaust nozzle, for example, during afterburner operation which can cause failure by burning through the nozzle wall. Various liner devices for the nozzles have been developed to treat this problem but difficulties in the manufacture and/or operation exist.
For example, one prior liner device placed a plurality of perforated panels in the nozzle. The panels are held and positioned away from the nozzle wall by means of hinged supports thus creating a cooling air plenum between the nozzle wall and the panels. Under extreme temperature operating conditions, cooling air is forced into the plenum and out of the perforated holes in the panels thus lowering the temperatures at the walls. The panels themselves move on the hinged supports in response to changes in pressure and temperature. This movement creates numerous problems because the various openings, non-cooling, must be sealed as best possible; otherwise, hot spots can develop on the nozzle wall. This particular liner device is more suitable to 2-D nozzles having flat walls; otherwise, the placement in a curved nozzle is difficult to accomplish.
Thus, there exists a need for an improved liner for a gas turbine nozzle that eliminates or substantially reduces sealing problems, panel movement problems, mounting problems on curved surfaces and can be easily attached to the substructure.